The present invention relates to spacecraft navigation, and, more particularly, to a system and method providing for on-board alignment of navigational accelerometers.
Spacecraft navigation is facing increased demands. Formerly, spacecraft navigation involved attaining a desired orbit and making occasional corrections to maintain the orbit. More recently, however, multiple-year missions involve, for example, more distant objectives of autonomous orbit change or re-entry. These subsequent missions require navigational systems capable of meeting objectives far more complex than maintaining orbit and far more enduring than navigational systems designed for attaining orbit.
One challenge faced in developing such an enduring navigational system is maintaining its calibration or alignment with respect to an inertial frame of reference. This is particularly true where it is impractical to perform this alignment on a regular basis from a ground or other remote station.
One navigational approach has been to mount three-axis accelerometers on a gyro-stabilized platform mounted on a launch vehicle or spacecraft. The gyro-stabilized three-axis accelerometers provide an acceleration history which can be integrated once to obtain a velocity history and integrated twice to obtain a position history for the spacecraft in a known coordinate system. The histories so obtained are used to attain elaborate navigational objectives. The alignment of the accelerometer is maintained by the characteristics of the gyro. Such an arrangement has been used effectively in guiding the ascent of a satellite to a predetermined orbit.
The mechanical limitations of such an arrangement in the face of launch stresses, environmental extremes in space, mechanical creep, other stresses and uncompensatable random drift rates and biases limit the long-term accuracy of the gyros of such an arrangement. Accordingly, the accelerometer-on-a-gyro approach has not proved satisfactory when a regenerative mission objective requires elaborate navigation months or years after orbit is first attained. Thus, such systems require remote realignment of accelerometers if such realignment is provided for at all. Other systems require dedicated celestial body trackers physically tied to the gyro platform to achieve the realignment and calibration of the gyros. Still other systems discuss a "strap-down" method as speculated in "Theory of Inertial Guidance" by Connie L. McClure, Prentice Hall 1960, pp. 286-291.
Spinning spacecraft have more enduring orientation aids such as attitude and nutation sensors. However, while these are well-adapted for maintaining and adjusting attitude, they are not suited for extensive spacecraft navigation. In other words, they cannot readily yield a substitute for the instantaneous acceleration histories provided by accelerometers during periods of thrust.
What is needed is a system and method for realignment of a navigational inertial coordinate system. The system should provide for self-alignment and operate for extended periods essentially autonomously, without the constraints of navigation gyros. The system should be capable of managing demanding navigational objectives after an extended orbital period.